Energy-efficient electro-thermal and electro-mechanical ice-protection method

ABSTRACT

A region of an airfoil to be protected from accreted ice is heated to reduce the strength of the bond between the ice and the airfoil, and is mechanically deformed to shed the accreted ice after the bond has been sufficiently weakened. Heating ceases before substantial water runback is generated. The mechanical deformation and the cessation of heating occur approximately simultaneously.

BACKGROUND OF THE INVENTION

The invention relates to ice protection, and more particularly relatesto ice protection systems for use on aircraft. In its most immediatesense, the invention relates to aircraft ice protection systems of thetype in which a semi-rigid skin forms a lifting surface of the wing andin which ice is removed by flexing the skin.

Ice contamination of lifting surfaces (e.g. wings and horizontal tails)is always disadvantageous because it interferes with airflow over thesurface. This in turn increases drag, reduces lift, and reduces theangle of attack at which the airfoil enters a stall. For this reason,airplanes are provided with systems that protect lifting surfaces fromexcessive levels of ice contamination on critical regions of the liftingsurface.

Ice protection systems vary widely in performance. This is becausedifferent types of aircraft lifting surfaces have differentsensitivities to ice contamination and, consequently, different iceprotection requirements. For example, certain lifting surfaces may useairfoils that are more or less tolerant to the effects of icecontamination than are other airfoils; a quantity of accreted ice thatmight only imperceptibly degrade the performance of one type of airfoilsection might be a severe hazard to another type of airfoil whenoperated at similar or different flight conditions.

Accordingly, different types of ice protection systems are optimizeddifferently; energy available for ice protection is selectively appliedto address the particular sensitivity of one airfoil section as opposedto another. On some airplanes, there may be sufficient electrical powerto operate an ice protection system that relies onelectrically-generated heat (e.g., an electro-thermal deicing oranti-icing system) while on others, the available electrical power isinsufficient for such a system. Such considerations necessarily affectthe selection of the type of ice protection system to be used. Forexample, on an airplane having ample electrical power or bleed air, anenergy intensive evaporative anti-icing system may be employed, whilefor a power-limited application a deicing system may be used to shed iceonly when the ice accumulation reaches a predetermined distributedthickness that has been shown to degrade the performance of the airfoilto an unacceptable extent.

Although many different types of ice protection systems are available toaddress a wide variety of applications for ice protection, oneapplication category is particularly problematic. This is when theairfoil is very sensitive to ice contamination and there is limitedpower available for operation of an ice protection system that meets therequired performance.

The present invention is suitable for applications of this type. Testshave demonstrated that a preferred embodiment of this invention canmaintain worst-case distributed ice accretions to within criticallimitations (typically less than 0.050 inch) while consuming only afraction of the power that would conventionally be expected to berequired for such an application.

The invention proceeds from a realization that an existing iceprotection system can be reconfigured to operate in an entirelydifferent and highly advantageous manner. Commonly-owned U.S. Pat. No.5,921,502 (incorporated herein by reference, and referred to hereinafteras the “'502 patent”) discloses a hybrid ice protection system in whichan airfoil has a semi-rigid skin 58. The skin 58 is caused to flex byactuators 50, 52, 54, and 56, and the leading edge region 4′ is heatedby an electrical heater 10′. In operation, the electrical heater 10′ isoperated as a running-wet anti-icer and ice accreted aft of the leadingedge region 4′—so-called “runback refreeze” ice—is periodically removedby the actuators 50, 52, 54, and 56. In essence, the '502 patentdiscloses an ice protection system that uses heat to prevent icecontamination where the airfoil is most roughness-sensitive and usesmechanical flexing of a semi-rigid skin to shed accreted ice fromlocations where the airfoil is less so.

The inventor of the present invention realized that energy efficiency ofsuch a hybrid system would be much improved if the heat and the flexingof the semi-rigid skin were employed together in a coordinated fashioninstead being used independently in different locations. In accordancewith the invention, heat and mechanical deformation are both applied toa region to be protected, but the heat is used only to increase thetemperature of the ice/skin interface, reducing the adhesion forcebetween the ice layer and the subjacent skin and thereby weakening thebond between the skin and the ice that has accreted upon it instead ofremoving the ice by melting it into water (as in the prior art). Oncethis bond has been weakened, two things happen approximatelysimultaneously: the heat is turned off, and the actuators are fired toflex the skin. Because the bond between the ice and the skin has beensubstantially reduced or eliminated, the flexing of the semi-rigid skincompletely sheds the accreted ice. Furthermore, the absence of heatingbetween deicing cycles causes the temperature of the skin to drop belowfreezing before runback refreeze can be created. Consequently, much lessoverall heat is delivered to the protected region, and this greatlyimproves the energy efficiency of the system.

Advantageously although not necessarily, and in the preferredembodiment, energy consumption of the system is further reduced byremoving accreted ice from the airfoil on a zone-by-zone basis. This isaccomplished by dividing the airfoil into a series of zones that extendalong the span of the airfoil and then applying heat to individual zonessequentially, one at a time. This avoids the power drain that would berequired to heat the entire airfoil at once. In particular, it ispossible to eliminate the continuously heated parting strip along theentire span of a typical electro-thermal de-icing system.

In the preferred embodiment, heating continues until a very thin layerof ice at the ice/skin interface is melted immediately adjacent theskin. This insures that the accreted ice is completely shed when theskin is deformed. However, this is only preferred, and it may bepossible to obtain acceptable performance even if the ice is notentirely melted at the surface of the skin.

In the preferred embodiment, the heating and flexing are co-located inthe region of the airfoil that is most sensitive to the effects ofaccreted ice. This applies the maximum heat and the maximum mechanicalforce to the accreted ice in the location where ice will most seriouslydegrade aerodynamic performance. However, it will be understood thateven if the ice to be removed is at some distance from the heater, theactuators, or both, it is nonetheless possible that the ice can beremoved. Furthermore, in accordance with the preferred embodiment,regions of the airfoil that are less sensitive to ice contamination areprotected only by flexing of the skin, as in the '502 patent.

In the preferred embodiment, heating is electro-thermal. However, thisis only preferred, and it may be possible to obtain acceptableperformance using bleed air as a source of heat.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be better understood with reference to the followingillustrative and non-limiting drawings, in which:

FIG. 1 is a flowchart schematically illustrating a method in accordancewith a preferred embodiment of the invention;

FIGS. 2 and 3 show apparatus in accordance with a preferred embodimentof the invention; and

FIGS. 4, 5A, 5B, and 5C schematically illustrate deicing of the wings ofan aircraft in accordance with a preferred embodiment of the invention;and

FIG. 6 is a schematic illustration of the system architecture of theelectronic circuitry used in apparatus in accordance with a preferredembodiment of the invention.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

The same element is always indicated by the same reference numeral inall views. The drawings are not necessarily to scale, and parts may beenlarged for clarity.

In the following description, the ice-protected component isspecifically illustrated to be an aircraft wing. This is because thepresent invention was designed for this application. However, theinvention is not limited to use on aircraft wings and can be used onother surfaces such as horizontal stabilizers, vertical fins, aircraftinlets, and other airfoils.

The following description also assumes that icing wind tunnel tests havebeen or will be carried out on a model that simulates, underdesigned-for conditions of flight, the accretion of ice on the componentto be protected. Such tests are routinely conducted during the design ofan aeronautical ice-protection system, because empirical data isnecessary to verify that the system performs as required within theentire performance envelope of the aircraft. Such tests will revealanomalies (e.g. locations on the airfoil where the heat required toweaken the bond between the skin and the accreted ice is either greaterthan or less than expected) and such anomalies will be corrected byappropriate decreases or increases in heat delivered to such locations.Thus, the following description necessarily describes the preferredembodiment of the invention in general terms; it is not possible to setout specific design details on an a priori basis.

I. Method in Accordance with a Preferred Embodiment

A. Theory of Operation

In accordance with the preferred embodiment of the invention, deicing isaccomplished using a unique combination of heat and flexing of theice-bearing surface. As is specifically discussed in the '502 patent,use of both heat and flexure in ice-protection systems of the semi-rigidskin type is already known.

However, in accordance with the preferred embodiment, heat and flexingare used in a specific manner that will be described in connection withFIG. 1. It will be understood that FIG. 1 relates to a single locationon the leading edge of an airfoil; this is to illustrate the principleof deicing that is implemented in the preferred embodiment of theinvention. The deicing of the airfoil as a whole, and the mechanism bywhich this is accomplished, are discussed afterward.

In accordance with the preferred embodiment, heat and flexing of theskin are applied to the same region of the airfoil and not (as isdisclosed in the '502 patent) to different regions. In an initial step100, ice has begun to accrete, and deicing is initiated. This may bedone at the command of the pilot or other crew member, or alternativelyautomatically by using an ice sensor. Upon initiation of deicing, theprotected region is heated in step 110 by energizing an electricalheater located there. (An electrical heater is preferred because it iseasy to control. But, it may alternatively be possible to use anotherheat source such as bleed air.) As is known to persons skilled in theart, when ice accretes on a surface and the temperature of theice-surface interface increases from cold towards 32° F., the adhesionforce between the ice and the surface decreases substantially. Thus, asthe electrical heater continues to operate and the temperature of theinterface increases, the accreted ice becomes easier to dislodge.

Heating continues until step 120, when the temperature of the protectedregion has been raised to a predetermined value at which the adhesionforce has been sufficiently reduced to permit adequate performance. (Intests, the assignee used a temperature of 42° F., but in practice thetemperature will vary with the location of the region to be deiced. Theassignee believes this temperature can be as high as 52° F. or as low as37° F., depending on the particular application. In practice, locationson the airfoil would be mapped to calculated temperatures and icing windtunnel tests carried out to identify locations in which the calculatedtemperatures were either excessive or insufficient.) In accordance withthe preferred embodiment of the invention, at approximately this timethe heat is turned off and then the ice-bearing surface is flexed (step130). Flexure of the ice-bearing surface sheds the cap of accreted icebecause the increase in the interface temperature has reduced thestrength of the bond between the ice cap and the surface.

Optionally, heater energization may continue for a short time (e.g. twoseconds) after the ice-bearing surface is flexed. This may beadvantageous at very cold ambient temperatures, to insure that themelted ice remains debonded and does not refreeze while the accreted icecap is being removed by aerodynamic forces. Alternatively, at relativelywarm ambient temperatures the ice-bearing surface may be flexed shortly(e.g., two seconds) after the heater has been turned off, because atsuch temperatures the ice at the interface with the skin does notrefreeze immediately. However, while there are conditions that may makeit advantageous for the heater to be turned off slightly before orslightly after the ice-bearing surface is flexed, it is important thatsuch flexure neither substantially precede, nor substantially follow,de-energization of the heater. If flexure comes too soon or too late,the ice may not be fully removed, and there may be places whereaccumulated ice exceeds the allowable ice accretion limit.

After the ice-bearing surface has been flexed, in step 140 the decisionis made whether to continue deicing. As in step 100, this decision canbe made automatically or by a crew member. If deicing is to bediscontinued, the deicing system is shut off. If deicing is to becontinued, then in step 150 the system waits until a determined state ofice accretion exists on the airfoil or for an appropriate cycle time andthe cycle repeats once again with step 110.

B. Preferred Control Methodology

Step 120 can be carried out in at least two ways. It is possible to puta temperature sensor at the protected region and to use the actualtemperature as a trigger to turn the heat off. Alternatively, it ispossible to avoid the complexity of a temperature sensor and associatedcontrol circuitry by calculating the duration of heating in advance andturning the heat off after the heater has been turned on for theappropriate duration. As stated above, this duration will typically beobtained from actual experimental data acquired during development testsin icing wind tunnels. It is specific to the surface geometry of theairfoil, true airspeed (TAS), and outside air temperature (OAT).

The appropriate duration of heating may be approximated by monitoringthe TAS of the aircraft and OAT, and applying semi-empirical equationsusing these variables in the calculation. Persons skilled in the artknow that the maximum quantity of liquid water content (LWC) in the airis a function of the OAT (the Federal Aviation Regulations, or FARs,establish design criteria that must be met in this respect). Suchpersons also know that the rate at which water-bearing air is incidenton the airfoil's impingement zone is a function of the aircraft's TAS.

The worst-case rate of ice accretion at any particular TAS and OAT cantherefore be computed, and from this it is possible to compute thequantity of energy required to raise the protected region to apredetermined temperature at which the adhesion between the ice and theice-accreting surface has been sufficiently reduced. Furthermore, thethermal power density produced by the heater is also known. (Inexperiments, the assignee used a heater that delivered power densitiesvarying from about 20 watts/in² to about 30 watts/in² along the chord ofthe airfoil at a tunnel test speed of 160 knots.) From the quantity ofheat required at any particular combination of OAT and TAS and the rateat which heat is delivered per unit time, it is possible to compute theduration of heating required to transfer the necessary quantity ofenergy for the worst case of ice accretion. The assignee has done thisusing a transient thermal computational model that simulates theaircraft heated skin with all the appropriate construction materials.

The assignee has verified the results from this model in an icing windtunnel, and calculated the durations in the table below for a TAS of 250knots (a common cruise/hold design speed). Assuming that the maximumacceptable ice accretion is 0.050 inch, and a maximum ice collectionefficiency of near 30%, the model produces the following results:

Liquid Total Water Heater On Temperature Content Cycle Time Time OAT (°F.) @250 kts (g/m³) (seconds) (seconds) 17.2 32.0 0.47 Near 30 1 14 28.80.43 33 1.4 −4 10.8 0.21 66 3.2 −22 −7.2 0.14 90 5.0(The −22° F. lower limit of the above table was chosen to accord withthe icing envelope defined in FAR 25, Appendix C. The total temperaturecolumn includes the effects of aerodynamic heating. At an OAT of 17.2°F., the actual temperature of accreted ice will be 32.0° F., so only asmall quantity of heat will be required to melt the ice to water.) Inall instances, the computed cycle time is about half the durationrequired for ice to accumulate to the maximum acceptable level. Thisprovides an additional margin of safety, and is not necessary.

In other words, at a TAS of 250 knots and an OAT of 17.2° F., the heaterwill be energized for 1.0 second, and during this time the temperatureat the interface between the accreted ice cap and the skin will riseresulting in an appropriately low adhesion strength. And, at or aboutthe time the heater is de-energized, the protected region is flexed,causing the ice cap to be shed. The heater remains de-energized for theremainder (approximately 29 seconds) of the cycle time, after which theheater is energized once more and the cycle begins again.

At a TAS of 250 knots and an OAT of −22° F., the heater will beenergized for 5.0 seconds, during which time the adhesive bond betweenthe accreted ice cap and the skin will be appropriately degraded. Theheater must remain on for a longer time than in the previous examplebecause the temperature of the accreted ice in this example must beraised by 49.2° F. to reach the set point temperature of 42° F. (in theprior example, the temperature of the accreted ice had to be raised byapproximately 10° F. to reach the 42° F. set point temperature). As inthe prior example, at or about the time the heater is de-energized, theprotected region is flexed, causing the ice cap to be shed. However, inthis instance the heater remains de-energized for a period of 85 seconds(90 seconds cycle time less 5 seconds heating time), instead of 29seconds as in the previous example. The reason cycle time is longer foran OAT of −22° F. than for OAT of 17.2° F. is because LWC at −22° F. islower than at 17.2° F., resulting in a reduced rate of ice accumulation.

In the above examples, the cycle time of step 150 is calculated from theOAT and the TAS. However, this is not required. It is alternativelypossible to use a properly positioned and calibrated ice rate sensor tovary the cycle time of step 150 in accordance with the ice accretionrate.

As stated above, although there are conditions under which the heatermay advantageously be turned off slightly before or after theice-bearing surface is flexed, it is nonetheless true that these twoevents must always occur close together in time. The objective is toflex the specified protected region at the proper time and sequence inorder to remove ice accretions sufficiently so that if any inter-cycleor accumulated ice remains, it will be within acceptable limits.

II. Apparatus in Accordance with a Preferred Embodiment

Referring now to FIGS. 2 and 3, an airfoil generally indicated byreference numeral 10 (in this example, the airfoil is an airplane wing)contains a rigid substructure generally indicated by reference numeral12. The function of the substructure is to prevent the airfoil 10 fromdeforming in use.

As is disclosed in the '502 patent, the breeze surface of the airfoil 10is made of a semi-rigid skin 14. The skin 14 is advantageously made of arelatively thin metal with a relatively high thermal diffusivity (highthermal conductivity and low thermal capacitance) for fast thermalresponse from the heater input, but it can alternatively be made ofanother metal or a fiber composite. The material and precise thicknessof the skin 14 are not features of the invention; it is only importantthat the skin 14 be sufficiently rigid to return to its unflexed stateeven after many flexing cycles while being sufficiently flexible to beflexed by actuators driven by relatively low-power electrical pulses(see below).

Aft of the protected region 16 (which, see below, is within or near theimpingement zone of the wing 10) the skin 14 may be thicker and is fixedto ribs 18 (which are part of the substructure 12). Thus, the skin 14 isunsupported at the protected region 16. A flexible and relatively thinelectric heater 17 (advantageously but not necessarily made of heaterwires, printed circuits or electrically conductive sheets embedded in ahigh thermal diffusivity carrier that is non-electrically conductive,e.g., a glass fiber composite or polyimide film) is co-located directlyon the inside surface of the skin 14 at the protected region 16.

Inside the wing, and aft of the protected region 16, is located a framestructure 20 (which, like the ribs 18, is part of the substructure 12).Pairs 22A, 22B etc. of actuators generally indicated by referencenumeral 22 in FIG. 2 are secured to the frame structure 20 as byfasteners 24, but the method of attaching the actuators 22 to the framestructure 20 is not part of the invention. The actuators 22 bear againstthe heater 17. All the actuators 22 are identical; as will be discussedbelow, the actuators 22A are advantageously fired as a pair, as are theactuators 22B, and as are other pairs of actuators (not shown).

The actuators 22 are of the type disclosed in U.S. Pat. No. 5,782,435and the above-referenced '502 patent. As is explained in those patents,each of the actuators 22 is in the shape of a flattened and elongatedtube. When an actuator 22 is fired by a short, high voltage electricalpulse, magnetic fields inside the actuator 22 momentarily change itsshape and the actuator 22 becomes more circular for a short period oftime. The frame structure 20 is rigid, the skin 14 and attached heater17 form a semi-rigid structure, and the actuators 22 bear against theheater 17. So, when the actuators 22 are pulsed, they momentarily stressthe skin 14 (along with the heater 17 that is attached to it) and forcethe skin 14 outward, causing it to flex momentarily to shed accreted iceas discussed above. Once the electrical pulse has ceased, the internalmagnetic fields inside the actuator 22 collapse, the actuator 22 returnsto its flattened state, the stress on the skin 14 and attached heater 17is removed, and the skin 14 (together with the attached heater 17)returns to its original position.

In the preferred embodiment, two actuators 22 are operated sequentiallyas a pair (generally, 3 hits per actuator per cycle, lasting about 0.3second per hit). This is to insure that the flexure of the skin 14 isadequate to shed the accreted ice. However, this is not required, and itis alternatively possible to use one actuator 22 or more than two ofthem. The number of actuators 22 will be determined by the requirementsof the application in which the invention is to be used.

III. Application to a Commercial Aircraft

For most efficient operation, apparatus in accordance with the inventionmust be appropriately positioned on the surface to be protected. Properpositioning will now be discussed with reference to FIG. 4, whichillustrates a cross-sectional view of the leading edge region of anairfoil 10, which (as discussed above) in this example is an airplanewing.

In icing conditions, droplets of supercooled water (not specificallyshown) impinge on the airfoil 10. Relative to the airfoil 10, thedroplets follow trajectories 300. As can be seen from FIG. 4, thedroplets do not impinge on the entire surface of the airfoil 10; theyimpinge only on the impingement zone 306 between the upper impingementlimit 302 and the lower impingement limit 304.

It is evident that any ice protection system must protect at least aportion of the impingement zone 306, because ice will surely accretethere and the forward portion of the impingement region is usually themost contamination-sensitive region of the airfoil. However, as apractical matter an ice protection system cannot be limited to theimpingement zone 306. This is because ice forming aft of the impingementzone 306 will—perhaps to an unacceptable extent—disturb the airflowrequired for efficient operation. For example, if the aircraft undergoesan icing encounter with supercooled large droplets, including freezingrain or drizzle, ice can contaminate the surface of the airfoil 10 farbeyond the periphery of the impingement zone 306. Furthermore, as statedabove, some ice protection systems generate runback water, which willgenerally refreeze aft of the impingement zone 306 to form ice ridges.These ice ridges can create a substantial discontinuity and cansubstantially change the shape of the breeze surface of the airfoil 10.

For these reasons, an ice protection system must in most cases preventexcessive ice accumulation throughout a larger region, shown in FIG. 4as the roughness-sensitive region 308. Hence, in accordance with thepreferred embodiment, the skin 14 must be attached to the airfoil aft ofthe roughness-sensitive region 308 so as to be able to flex within allparts of it.

In accordance with the preferred embodiment, the heater 17 is locatedonly in the maximally contamination-sensitive region of the airfoil,i.e. in the region where accreted ice must be held within stringentlythin limits. (This region is identified by the airframe manufacturer.)As stated above, in this maximally contamination-sensitive region,accreted ice (not shown) is removed by co-action between heat from theheater 17 and flexure of the skin 14; heat is used to reduce the forceof adhesion between the ice and the skin 14, and flexure of the skin 14then removes the ice. Region 310, which is located within theroughness-sensitive region 308 but is outside the periphery of theheater 17, is protected only by flexure of the skin 14. This is becauseregion 310 is less sensitive to ice contamination than the regioncovered by the heater 17, and ice accumulation there need not be held tosuch stringently thin limits. The actuators used in the preferredembodiment are consistently able to remove ice layers that are 0.060inch or more thick. In the illustrated example, the airfoil is assumedto have acceptable performance even when ice contamination in region 310reaches a thickness of 0.060 inch. Thus, there is no need to provideheat to region 310 and a heater is unnecessary there.

In practice, the airframe manufacturer will define the overallroughness-sensitive region 308 and will identify the maximallycontamination-sensitive region where the heater 17 is to be located. Theless contamination-sensitive region 310 will then be defined by default.It is alternatively possible (although unlikely) for more than onemaximally contamination sensitive region to exist, and if this is so aplurality of heaters 17 can be used and controlled separately ortogether as the application requires. In other words, in accordance withthe preferred embodiment there is at least one maximallycontamination-sensitive region; there may be more than one in particularcircumstances. Likewise, in accordance with the preferred embodimentthere will usually be a less contamination-sensitive region in which themaximum acceptable ice accumulation thickness is greater, but this isnot absolutely necessary. There may exist an ultra-high performanceairfoil wherein the entire roughness-sensitive region must be kept veryclean under all icing conditions, and for such a demanding applicationthe heater 17 would be precisely coextensive with theroughness-sensitive region 308.

FIGS. 5A-5C shows how a method in accordance with the preferredembodiment of the invention can be implemented on the wings of aconventional commercial jet. Each wing is provided with four spanwiseslats, SL1 through SL4 on the left wing and SR1 through SR4 on the rightwing, and each spanwise slat (e.g. SR1) is provided with nine protectedzones Z1SR1, Z2SR1 . . . Z9SR1.

Apparatus in accordance with the preferred embodiment of the inventionillustrated in FIGS. 2 and 3 is installed in each of the seventy twoprotected zones. (There are two wings, each with four slats, and eachslat has nine zones.) Four energy storage bank units (“ESBs”) 500, 502,504, and 506 are provided. ESB 500 provides energy to the apparatus inslats SL1 and SL2, ESB 502 provides energy to the apparatus in slats SL3and SL4, and ESBs 504 and 506 similarly provide energy to apparatus inslats SR1 and SR2 and slats SR3 and SR4, respectively. As is discussedbelow, each ESB includes the functionality necessary to turn heaters 17on and off and to fire actuator pairs—e.g. actuators 22A and 22B—insuccession. A deicing control unit (“DCU”) 508 is connected to theavionics system 510 of the aircraft, which supplies information aboutthe aircraft's TAS and OAT to the DCU so that it can appropriatelycontrol the operation of the ESBs 500-506.

In operation, when a deicing operation is initiated by the DCU 508, eachESB 500, 502, 504, 506 operates identically. In the instance of ESB 504(see FIGS. 5B and 5C), the heater 17 in the first zone Z1SR1 of thefirst slat SR1 is turned on, the bond between the accreted ice and thefirst zone Z1SR1 is reduced, the heater 17 is turned off, and theactuators 22 in the first zone Z1SR1 are fired to remove the accretedice. Then, the same operations are carried out in the second zone Z2SR1.This process then continues, zone by zone progressing from inboard tooutboard, until the last (ninth) zone Z9SR1 has been cleared of ice.

Once all the zones Z1SR1-Z9SR1 have been cleared of ice, the ESB 504then repeats this operation with the nine zones Z1SR2-Z9SR2 in thesecond slat SR2, clearing one zone after the next, progressing frominboard to outboard. Once this has been completed, the ESB 504 thenrepeats the same operation in the first slat SR1.

It will be understood that all the ESB's 500-506 operate in parallel, sothat at any given time four of the eight slats SR1-SR4 and SL1-SL4 willbe undergoing deicing. In this illustrated preferred embodiment, slatsSR1, SR3, SL1, and SL3 are deiced together, and then slats SR2, SR4,SL2, and SL4 are deiced together. And, in each instance, one zone ineach slat is deiced at a time, with the most inboard zone being deicedfirst and the most outboard zone being deiced last. The whole cycle isrepeated as necessary, at such speed as is required to maintain thewings at or below the maximally-acceptable levels of ice accretion.

Persons skilled in the art will realize that the particular arrangementof slats and zones is not critical. Other arrangements of slats andzones, and other orders of operation (e.g. outboard to inboard asopposed to inboard to outboard, or a nondirectional order) can be usedinstead. Furthermore, in the preferred embodiment four ESB's 500-506 areused, each supplying energy to two slats (e.g. SR1 and SR2). Thisarrangement is preferred because it minimizes the weight of the systemand keeps the distances between an ESB and the zones it deices to aminimum, thereby maximizing energy transfer to the actuators andminimizing energy loss in the wiring. However, this is not required. OneESB may be used for each slat, or for more than two slats, or even forall zones on a single wing, depending on the application and theparticular components chosen for the ESB.

In this connection, one important additional consideration indetermining the number of ESB's in the system is the desirability ofavoiding wide fluctuations in the power drawn from the aircraft whilestill preventing ice accretion from exceeding the preestablished maximumacceptable thicknesses. In this example, only two ESB's (e.g. theinboard ESB's 500 and 504 or the outboard ESB's 502 and 506) operatesimultaneously. As compared with using four ESB's, this doubles theduration of the cycle required to deice all seventy two zones, butreduces by half the power required to operate the system. Such anarrangement prevents unnecessarily wide swings in the power drawn by thesystem, i.e. avoids high power draw while deicing is ongoing followed bynegligible power draw during a long idle period following a completedeicing cycle.

Additionally, while the slats SR1-SR4 and SL1-SL4 are shown to beapproximately the same length, and the various zones (e.g. Z1SR1-Z9SR1)are shown as being approximately the same width, this is merely aschematic. The dimensions of the various components will be dictated bythe particular application for which the system is used.

FIG. 6 shows in more detail the system architecture of the electricalelements of apparatus in accordance with the preferred embodiment. EachESB actually contains three main component subgroups. One subgroup is abank of capacitors or other energy storage devices, together withcircuitry that keeps the capacitors at an appropriate state of chargeduring operation. Another subgroup is made up of components that causevoltage pulses from the capacitors to be directed to the particularactuators or actuator pairs 22 to be fired. The third subgroup is madeup of components that cause the heaters 17 to be turned on and off. Inoperation, the DCU 508 initiates deicing when an ice detector (notshown) sends a signal (ICE DETECTOR) to the DCU 508 that ice isbeginning to accumulate. The DCU supplies power to the heaters 17(HEATER POWER) through the ESB's, and high voltage (HIGH VOLTAGE) tocharge the capacitors in the ESB's 500-506. Then, in accordance with OATand TAS information (OAT and TAS) from the aircraft's avionics system510, the DCU 508 sends out timed trigger signals (TRIGGER) that causethe heaters 17 and pairs of actuators 22 to be properly energized andde-energized.

In this preferred embodiment, the DCU 508 will accept input from theflaps (FLAPS STATUS). This is to provide the option of providing anextra-clean wing surface on final approach (when the flaps will beused). This is particularly true when conditions are such that a pauseperiod exists following the completion of one deicing cycle and beforethe commencement of the next one. In such a case, the pause period maybe eliminated and the deicing cycle shortened, thus decreasing themaximum accumulation of inter-cycle ice. The DCU provides statusinformation (COCKPIT INDICATORS) to the pilot.

Optionally, and in accordance with the preferred embodiment, atemperature sensor 512 is provided for each of the ESB's 500-506. Thetemperature sensor 512 is mounted to the skin 14 or the heater 17 of thefirst zone deiced by each ESB (e.g. zone Z1SR1 in the case of ESB 504)and measures the temperature when the heater 17 is energized. This canbe used to make sure that the skin 14 is brought up to the propertemperature when the heater 17 is energized, and to provide a means toadjust the duty cycles of the heaters 17 if this is not so. Fourtemperature sensors 512 provide redundancy. When the sensor is installedon the skin, it provides a direct status of the skin temperature.However, when it is installed on the heater, it provides an indirectstatus of the skin temperature, which is calculated based on the heatersensor temperature, sensor location, heater construction, materialproperties, and heater power density (besides TAS and OAT). Clearly,installation of the sensor on the skin is preferred, but is notnecessarily the only practical alternative.

Quite obviously, the herein-disclosed system architecture is preferredbut not required. It would alternatively be possible to provide morethan one DCU 508, and the particular system inputs and outputs to theDCU 508 and the ESB's 500-506 could be different. The systemarchitecture is not part of the invention.

Although at least one preferred embodiment of the invention has beendescribed above, this description is not limiting and is only exemplary.The scope of the invention is defined only by the claims, which follow:

1. (canceled)
 2. (canceled)
 3. (canceled)
 4. (canceled)
 5. (canceled) 6.(canceled)
 7. A method of deicing a protected region of an airfoilsurface, comprising the following steps: heating the protected region ofthe surface to a temperature that will melt ice at its interface withthe protected region; ceasing to heat the protected region of thesurface once said temperature has been reached; and approximatelysimultaneously with said ceasing step, mechanically deforming theprotected region of the surface to shed ice accreted thereon.
 8. Amethod of deicing a protected region of an airfoil surface, comprisingthe following steps: heating the protected region of the surface to atemperature that will acceptably reduce the force of adhesion of ice tothe protected region of the surface; ceasing to heat the protectedregion of the surface once said temperature has been reached; andapproximately simultaneously with said ceasing step, mechanicallydeforming the protected region of the surface to shed ice accretedthereon.
 9. A method of deicing a protected region of an airfoilsurface, comprising the following steps: electro-thermally heating theprotected region of the surface to a temperature that will acceptablyreduce the force of adhesion of to the protected region of the surface;ceasing to heat the protected region of the surface once saidtemperature has been reached; and approximately simultaneously with saidceasing step, mechanically deforming the protected region of the surfaceto shed ice accreted thereon.
 10. A method of deicing a protected regionof an airfoil surface, comprising the following steps: electro-thermallyheating the protected region of the surface to a temperature that willmelt ice at its interface with the airfoil; ceasing to heat theprotected region of the surface once said temperature has been reached;and approximately simultaneously with said ceasing step, mechanicallydeforming the protected region of the surface to shed ice accretedthereon.
 11. The method of claim 10, wherein said heating step isinitiated by a crew member.
 12. The method of claim 10, wherein saidheating step is initiated automatically.
 13. The method of claim 10,wherein said ceasing step is carried out automatically at the end of aperiod of time that is determined based upon true airspeed and outsideair temperature.
 14. The method of claim 10, wherein said ceasing stepis carried out automatically based upon actual temperature of theprotected region.
 15. A method of deicing a protected region of theaerodynamic surface of an airfoil in flight when the airfoil is movingat a known true airspeed (TAS) through air at a known outside airtemperature (OAT), comprising the following steps: heating the protectedregion of the aerodynamic surface for a period of time determined by theTAS and OAT; momentarily flexing the protected region of the aerodynamicsurface at approximately the end of said period of time; and repeatingsaid heating and flexing steps.
 16. The method of claim 15, wherein saidheating step is carried out using electro-thermal heating. 17.(canceled)
 18. (canceled)
 19. A method of deicing the protected regionof an airfoil in flight, comprising: heating the protected region of theairfoil to a predetermined temperature above freezing; and momentarilyflexing the protected region of the airfoil no later than when theprotected region reaches said predetermined temperature.
 20. The methodof claim 18, wherein the predetermined temperature is between 37° F. and52° F.